Space Mission Energy Management Architecture

ABSTRACT

A system operating in a spacecraft includes a tank storing a propellant, a thermal receiver configured to change the propellant from a first phase to a second phase by providing heat to the propellant, and an energy conversion device fluidically coupled to the tank and configured to generate electric energy using the propellant in the second phase.

CROSS-REFERENCE TO RELATED APPLICATIONS

This is a continuation of U.S. patent application Ser. No. 16/773,880,filed Jan. 27, 2020, which claims the benefit of U.S. Provisional PatentApplication No. 62/814,496, filed on Mar. 6, 2019, and U.S. ProvisionalPatent Application No. 62/956,091, filed on Dec. 31, 2019. Thedisclosure of each of the above-identified applications is incorporatedherein by reference in its entirety.

FIELD OF THE DISCLOSURE

The disclosure generally relates to operating a spacecraft and morespecifically to sharing energy and propellant resources among thrustersof different types.

BACKGROUND

With increased commercial and government activity in the near space, avariety of spacecraft are under development, for various missions. Forexample, some spacecraft may be dedicated to delivering payloads (e.g.,satellites) from one orbit to another. In such orbital transfermissions, the spacecraft generally requires a thruster that consumes apropellant in order to achieve mission objectives.

SUMMARY

This disclosure generally relates to improving efficiency of using apropellant and collected solar energy in a spacecraft system. Thepropellant and/or the energy may be shared between two thrusters,operating according to different propulsion techniques. The system mayuse the propellant to store, direct, or transform collected solarenergy.

In one aspect, a system operating in a spacecraft includes a tankstoring a propellant, a thermal receiver configured to change thepropellant from a first phase to a second phase by providing heat to thepropellant, and an energy conversion device fluidically coupled to thetank and configured to generate electric energy using the propellant inthe second phase.

In another aspect, a method implemented in a spacecraft includes storinga propellant in a tank, changing the propellant from a first phase to asecond phase by providing heat to the propellant via a thermal receiver,and generating, by an energy conversion device fluidically coupled tothe tank, electric energy using the propellant in the second phase.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a block diagram of an example spacecraft configured fortransferring a payload between orbits.

FIG. 2 illustrates an example configuration of a spacecraft system inwhich a controller controls supply of propellant to two thrusters.

FIG. 3 illustrates an example configuration of a spacecraft system inwhich a controller controls conversion, distribution, and/or consumptionby thrusters of a common supply of solar energy.

FIG. 4 illustrates an example configuration of a spacecraft system inwhich a controller is configured to use a solar concentrator to directradiant solar energy to an energy plant or to a solar-thermal thruster.

FIG. 5 illustrates an example configuration of a spacecraft system inwhich a controller 542 may direct propellant through one of the two heatexchangers within a thermal receiver.

FIG. 6 illustrates an example configuration of a spacecraft system whichconverts solar energy into propellant enthalpy, and uses a controller tocause the heated propellant to be directed to a thruster or to a turbineto generate electric energy.

FIG. 7 illustrates an example configuration of a spacecraft system whichuses propellant as a working fluid for a fuel cell.

FIG. 8 illustrates an example spacecraft with two thrusters moving alonga trajectory, where a controller is configured to activate one or moreof the two thrusters based at least in part on a position along thetrajectory.

FIG. 9 is a flow diagram of an example method of operating a propulsionsystem of a spacecraft.

DETAILED DESCRIPTION

A spacecraft of this disclosure may be configured for transferring apayload from a lower energy orbit to a higher energy orbit according toa set of mission parameters. The mission parameters may include, forexample, a time to complete the transfer and an amount of propellantand/or fuel available for the mission. Generally, the spacecraft maycollect solar energy and use the energy to power one or more thrusters.Different thruster types and/or operating modes may trade off the totalamount of thrust with the efficiency of thrust with respect to fuel orpropellant consumption, defined as a specific impulse.

The spacecraft in some implementations includes thrusters of differenttypes to improve the efficiency of using solar energy when increasingorbital energy. In some implementations, the spacecraft uses the samesubsystems for operating the different-type thrusters, thereby reducingthe mass and/or complexity of the spacecraft, and thus decreasingmission time while maintaining and/or improving reliability.Additionally or alternatively, the spacecraft can choose or alternatebetween thrusters of different types as primary thrusters. Thespacecraft can optimize these choices for various mission goals (e.g.,different payloads, different destination orbits) and/or missionconstraints (e.g., propellant availability). Example optimization ofthese choices can include variations in collecting and storing solarenergy as well as in controlling when the different thrusters use theenergy and/or propellant, as discussed below.

FIG. 1 is a block diagram of a spacecraft 100 configured fortransferring a payload between orbits. The spacecraft 100 includesseveral subsystems, units, or components disposed in or at a housing110. The subsystems of the spacecraft 100 may include sensors andcommunications components 120, mechanism control 130, propulsion control140, a flight computer 150, a docking system 160 (for attaching to alaunch vehicle 162, one or more payloads 164, a propellant depot 166,etc.), a power system 170, a thruster system 180 that includes a firstthruster 182 and a second thruster 184, and a propellant system 190.Furthermore, any combination of subsystems, units, or components of thespacecraft 100 involved in determining, generating, and/or supportingspacecraft propulsion (e.g., the mechanism control 130, the propulsioncontrol 140, the flight computer 150, the power system 170, the thrustersystem 180, and the propellant system 190) may be collectively referredto as a propulsion system of the spacecraft 100.

The sensors and communications components 120 may several sensors and/orsensor systems for navigation (e.g., imaging sensors, magnetometers,inertial motion units (IMUs), Global Positioning System (GPS) receivers,etc.), temperature, pressure, strain, radiation, and other environmentalsensors, as well as radio and/or optical communication devices tocommunicate, for example, with a ground station, and/or otherspacecraft. The sensors and communications components 120 may becommunicatively connected with the flight computer 150, for example, toprovide the flight computer 150 with signals indicative of informationabout spacecraft position and/or commands received from a groundstation.

The flight computer 150 may include one or more processors, a memoryunit, computer readable media, to process signals received from thesensors and communications components 120 and determine appropriateactions according to instructions loaded into the memory unit (e.g.,from the computer readable media). Generally, the flight computer 150may be implemented any suitable combination of processing hardware, thatmay include, for example, applications specific integrated circuits(ASICs) or field programmable gate arrays (FPGAs), and/or softwarecomponents. The flight computer 150 may generate control messages basedon the determined actions and communicate the control messages to themechanism control 130 and/or the propulsion control 140. For example,upon receiving signals indicative of a position of the spacecraft 100,the flight computer 150 may generate a control message to activate oneof the thrusters 182, 184 in the thruster system 180 and send themessage to the propulsion control 140. The flight computer 150 may alsogenerate messages to activate and direct sensors and communicationscomponents 120.

The docking system 160 may include a number of structures and mechanismsto attach the spacecraft 100 to a launch vehicle 162, one or morepayloads 164, and/or a propellant refueling depot 166. The dockingsystem 160 may be fluidically connected to the propellant system 190 toenable refilling the propellant from the propellant depot 166.Additionally or alternatively, in some implementations at least aportion of the propellant may be disposed on the launch vehicle 162 andoutside of the spacecraft 100 during launch. The fluidic connectionbetween the docking system 160 and the propellant system 190 may enabletransferring the propellant from the launch vehicle 162 to thespacecraft 100 upon delivering and prior to deploying the spacecraft 100in orbit.

The power system 170 may include components (discussed in the context ofFIGS. 4-7) for collecting solar energy, generating electricity and/orheat, storing electricity and/or heat, and delivering electricity and/orheat to the thruster system 180. To collect solar energy into the powersystem 170, solar panels with photovoltaic cells, solar collectors orconcentrators with mirrors and/or lenses, or a suitable combination ofdevices may collect solar energy. In the case of using photovoltaicdevices, the power system 170 may convert the solar energy intoelectricity and store it in energy storage devices (e.g, lithium ionbatteries, fuel cells, etc.) for later delivery to the thruster system180 and other spacecraft components. In some implementations, the powersystem 180 may deliver at least a portion of the generated electricitydirectly to the thruster system 180 and/or to other spacecraftcomponents. When using a solar concentrator, the power system 170 maydirect the concentrated (having increased irradiance) solar radiation tophotovoltaic solar cells to convert to electricity. In otherimplementations, the power system 170 may direct the concentrated solarenergy to a solar thermal receiver or simply, a thermal receiver, thatmay absorb the solar radiation to generate heat. The power system 170may use the generated heat to power a thruster directly, as discussed inmore detail below, to generate electricity using, for example, a turbineor another suitable technique (e.g., a Stirling engine). The powersystem 170 then may use the electricity directly for generating thrustor store electric energy as briefly described above, or in more detailbelow.

The thruster system 180 may include a number of thrusters and othercomponents configured to generate propulsion or thrust for thespacecraft 100. Thrusters may generally include main thrusters that areconfigured to substantially change speed of the spacecraft 100, or asattitude control thrusters that are configured to change direction ororientation of the spacecraft 100 without substantial changes in speed.In some implementations, the first thruster 182 and the second thruster184 may both be configured as main thrusters, with additional thrustersconfigured for attitude control. The first thruster 182 may operateaccording to a first propulsion technique, while the second thruster 184may operate according to a second propulsion technique.

For example, the first thruster 182 may be a microwave-electro-thermal(MET) thruster. In a MET thruster cavity, an injected amount ofpropellant may absorb energy from a microwave source (that may includeone or more oscillators) included in the thruster system 180 and, uponpartial ionization, further heat up, expand, and exit the MET thrustercavity through a nozzle, generating thrust.

The second thruster 184 may be a solar thermal thruster. In oneimplementation, propellant in a thruster cavity acts as the solarthermal receiver and, upon absorbing concentrated solar energy, heatsup, expands, and exits the nozzle generating thrust. In otherimplementations, the propellant may absorb heat before entering thecavity either as a part of the thermal target or in a heat exchange withthe thermal target or another suitable thermal mass thermally connectedto the thermal target. In some implementations, while the propellant mayabsorb heat before entering the thruster cavity, the thruster system 180may add more heat to the propellant within the cavity using anelectrical heater or directing a portion of solar radiation energy tothe cavity.

The propellant system 190 may store the propellant for use in thethruster system 180. The propellant may include water, hydrogenperoxide, hydrazine, ammonia or another suitable substance. Thepropellant may be stored on the spacecraft in solid, liquid, and/or gasphase. To that end, the propellant system 190 may include one or moretanks. To move the propellant within the spacecraft 100, and to deliverthe propellant to one of the thrusters, the propellant system mayinclude one or more pumps, valves, and pipes. As described below, thepropellant may also store heat and/or facilitate generating electricityfrom heat, and the propellant system 190 may be configured, accordingly,to supply propellant to the power system 170.

The mechanism control 130 may activate and control mechanisms in thedocking system 160 (e.g., for attaching and detaching payload orconnecting with an external propellant source), the power system 170(e.g., for deploying and aligning solar panels or solar concentrators),and/or the propellant system (e.g., for changing configuration of one ormore deployable propellant tanks). Furthermore, the mechanism control130 may coordinate interaction between subsystems, for example, bydeploying a tank in the propellant system 190 to receive propellant froman external source connected to the docking system 160.

The propulsion control 140 may coordinate the interaction between thethruster system 140 and the propellant system 190, for example, byactivating and controlling electrical components (e.g., a microwavesource) of the thruster system 140 and the flow of propellant suppliedto thrusters by the propellant system 190. Additionally oralternatively, the propulsion control 140 may direct the propellantthrough elements of the power system 170. For example, the propellantsystem 190 may direct the propellant to absorb the heat (e.g., at a heatexchanger) accumulated within the power system 170. Vaporized propellantmay then drive a power plant (e.g., a turbine, a Stirling engine, etc.)of the power system 170 to generate electricity. Additionally oralternatively, the propellant system 190 may direct some of thepropellant to charge a fuel cell within the power system 190.

The subsystems of the spacecraft may be merged or subdivided indifferent implementations. For example, a single control unit maycontrol mechanisms and propulsion. Alternatively, dedicated controllersmay be used for different mechanisms (e.g., a pivot system for a solarconcentrator), thrusters (e.g., a MET thruster), valves, etc. In thefollowing discussion, a controller may refer to any portion orcombination of the mechanism control 130 and/or propulsion control 140.

FIG. 2 illustrates a configuration of a spacecraft system 100 in which acontroller 242 controls supply of propellant to two thrusters 282, 284from a shared propellant system that includes two valves 292, 294 tocontrol flow rate of the propellant from a tank 296. The valve 292 maybe disposed in a fluid line connecting the tank 296 to the thruster 282;and the valve 284 may be disposed in a fluid line connecting the tank296 to the thruster 284. The controller 242, configured to operate thevalves 292, 294 to control the supply of the propellant to the thrusters282, 284, may be included in the propulsion control 140 of FIG. 1. Thethrusters 282, 284 may be, correspondingly, the thrusters 182, 184. Thethruster 282 may be a solar thermal thruster, a resistojet thruster, orany other suitable thruster. The thruster 284 may be, for example, a METthruster that uses energy from a microwave source 286 to ionize thepropellant. In some implementations, the propellant for at least one ofthe thrusters 282, 284 may flow from a dedicated accumulator tank thatis fluidically connected (e.g., with pipes, valves, and/or pumps) to thecommon propellant tank 296. In any case, the controller 242 may selectwhich thruster(s) consume the shared supply of propellant to generatethrust. In some implementations, the controller 242 may direct the flowof the propellant to generate electrical power for the microwave source286, as discussed below. Additionally or alternatively, the controller242 may direct the flow of the propellant to cool the operatingmicrowave source 286

The controller 242 may direct the propellant to the thrusters 282, 284to efficiently utilize the difference in available thrust and/orspecific impulse of each thruster in optimizing propellant consumptionand/or reducing mission time. For example, the thruster 282 may providea higher thrust than the thruster 284 at the expense of a reducedspecific impulse. The controller 242 may be configured to activate thethruster 282 based on a position within an orbital transfer maneuver,for example, to take advantage of the Oberth effect and maximize theefficiency of propellant consumption while reducing mission time. Thecontroller 242 may be configured to activate the thruster 284 at adistinct section of the orbital transfer maneuver than the sectioncorresponding to activating the thruster 282, as discussed in moredetail in the discussions corresponding to FIGS. 8 and 9.

FIG. 3 illustrates a configuration of a spacecraft system 100 in which acontroller 342 controls conversion, distribution, and/or consumption bythrusters 382, 384 of a common supply of power from a power system 370.The power system can 370 can provide solar power, for example. Thecontroller 342 may be included in the mechanism control 130 and/or thepropulsion control 140 of FIG. 1. The thrusters 282, 284 may be,correspondingly, the thrusters 182, 184. As described in more detailbelow, the controller 342 may direct at least a portion of the output ofthe power system 370 in radiant from to one of a number of thermalreceiver, or may direct the flow of heat or electric energy (e.g., to amicrowave source 386) generated by the power system 370. In any case,the controller 342 may select which thruster(s) consume the sharedsupply of solar energy to generate thrust.

FIG. 4 illustrates a configuration of a spacecraft system 100 in which acontroller 442 (e.g., the controller 342) is configured to direct, usinga solar concentrator 472 (configured as a component of the power system170), radiant solar energy to an energy plant 474 or to a solar-thermalthruster 482. In some implementations, the solar concentrator 472 mayinclude one or more secondary mirrors, lenses, and/or fiber-optic guidesthat the controller 442 cause to move to route the radiant energy. Thesolar concentrator 472 may be pivotable (i.e., attached to a mechanismwith a pivot), and the controller 442 may be configured to cause thesolar concentrator to pivot toward the energy plant 474 or the solarthermal thruster 482, for example. The solar thermal thruster 482 mayinclude a thermal receiver. The thermal receiver may include a solidheat exchanger that absorbs solar energy and transfers it to thepropellant fluid. Additionally or alternatively, the thermal receivermay include the propellant fluid directly absorbing the solar radiation.To that end, one or more additives (e.g., metal powders) may be includedin the propellant.

The energy plant 474 may be configured to convert solar energy intoelectricity and deliver the electricity to a thruster. The thruster maybe a MET thruster 484 and delivering electricity may include powering amicrowave source 486. In some implementations, the energy plant mayinclude photovoltaic devices. In other implementations, the energy plantmay include a thermal receiver for converting radiant energy into heatand using heat to generate electric energy (e.g., using a turbine drivenby heated propellant, a suitable heat engine, or thermoelectric devices)as described in the examples below. The energy plant 474 need not powerthe microwave source 486 directly nor continuously. The energy plant mayinclude a battery, a fuel cell, or any suitable combination of electricenergy storage devices that may supply power to the microwave source 486in accordance with a controller (e.g., the controller 442, propulsioncontrol 140 , etc.).

FIG. 5 illustrates a configuration of a spacecraft system 100 in which acontroller 542 (e.g., the controller 242) may direct propellant from thecommon tank 296 through a first heat exchanger 572 or a second heatexchanger 574, the two heat exchangers configured as a thermal receiverfor the solar concentrator 576. The heat exchanger 574 may befluidically connected to an inlet of a turbine 578. A propellant stream592 that absorbs thermal energy in the first heat exchanger 572 may usethe thermal energy in thruster 582. On the other hand, a propellantstream 594 that absorbs thermal energy in the second heat exchanger 574may generate electric energy for operating a MET thruster 584 bypowering a microwave source 586. To that end, the propellant stream 594may drive the turbine 578. In some implementations, the turbine 578 maybe replaced with a suitable heat engine (e.g., a Stirling engine).

The heat exchangers 572, 574 may be configured as a thermal receiverthat includes a surface that absorbs the radiative energy falling on thethermal receiver and a thermal mass that conducts the heat away from thesurface. The heat exchangers 572, 574 may include corresponding fluidicchannels guiding the propellant past the thermal mass. The thermal massof the thermal receiver/heat exchangers may include a salt (e.g., sodiumnitrate, calcium nitrate, etc.), a metal (e.g., aluminum, copper, iron,etc.) or another suitable material configured to melt and store thermalenergy in the liquid phase and/or phase transition.

In operations, the controller 542 may direct the propellant to thefluidic channel of the heat exchanger 572, with the fluidic channelguiding the propellant to the thruster 582. Additionally oralternatively, the controller 542 may direct the propellant to thefluidic channel of the heat exchanger 574, with the fluidic channelguiding the propellant to the turbine 578 for generating electricity forthe thruster 582. The outlet of the turbine 578 may be fluidicallyconnected to the tank 296 or, in some implementations, to the thruster582 without returning to the tank 296. Thus, the fluidic channel of theheat exchanger 574 may guide the propellant toward the thruster 584.

FIG. 6 illustrates a configuration of a spacecraft system 100 whichconverts solar energy into propellant enthalpy, and uses a controller642 to direct the heated propellant to a thruster 682 via a valve 692,or, via a valve 694, to a turbine 678 (or any suitable energy plant) togenerate electric energy. The thruster 682 may be a solar thermalthruster configured as one of the main thrusters, or, in someimplementations, as an attitude adjustment thruster. The power system170 may be configured to use the electric energy generated from theheated propellant to power a microwave source 686 for an MET thruster694. The MET thruster 694 may consume the propellant used for generatingelectric energy. To that end, the outlet of the turbine 678 may befluidically connected with the MET thruster 694 (e.g., via one or morepipes, valves, pumps, accumulator tanks, etc.). Additionally oralternatively, the propellant system may return at least some of thepropellant used to generate electric energy to the tank 696.

A propellant tank 696 may be configured to include a thermal receiver toconvert to heat the solar radiation, directed to the tank 696 by thesolar collector 672. The propellant may vaporize upon absorbing the heatfrom the thermal receiver, storing the collected solar energy. At leasta portion of the tank 696 may be insulated and/or reflective to reducethe loss of propellant heat to space. Additionally or alternatively, thetank 696 may be expandable. The tank 696 may expand, for example, tomaintain a suitable pressure range as the amount of heated propellant inthe tank 696 changes.

The turbines 578, 678 in FIGS. 5 and 6, respectively, (or other energyconversion devices) need not power the corresponding microwave sources586, 686 directly nor continuously. Instead the turbines 578, 678 (orother suitable power plants) may charge a battery, a fuel cell, or anysuitable combination of electric energy storage devices that may supplypower to the corresponding microwave sources 586, 686. An exampleelectric energy storage implementation is illustrated in FIG. 7.

FIG. 7 illustrates a configuration of a spacecraft system 100 which usespropellant as a working fluid for a fuel cell 775. A turbine 778 may beconfigured to supply electricity for cracking water (or another suitablefluid that may be used as a propellant for a MET thruster 784) tothereby charge the fuel cell 775. The fuel cell 775 may power amicrowave source 786 directly or, for example, by way of a power unit777. The power unit 777 may include circuits for power conditioning,and/or switching. In some implementations, the power unit 777 mayinclude additional energy storage (e.g., a battery) configured to supplypower to the generator 786 when the fuel cell 775 is discharged and theMET thruster 784 uses the remaining propellant in the system 100. Inother implementations, the MET thruster 784 is configured to use thepropellant (e.g., water) produced as a byproduct (e.g., of recombinationof hydrogen and oxygen) of the fuel cell 775 as the fuel cell 775 powersthe MET thruster.

FIG. 8 illustrates a spacecraft 800 (which may be the spacecraft 100)with thrusters 882, 884 moving along an orbit 810 (or any suitableportion of an orbital transfer maneuver trajectory). A controller may beconfigured to activate the thruster 882 of the spacecraft 800 at asegment 812 of the orbit 810 and/or the thruster 884 of the spacecraft800 at segments 814 a,b of the orbit 810. The segment 812 may be asegment substantially closest to the periapsis (closest to the planetand lowest potential energy point) of the orbit 810, while the segments814 a,b may be adjacent to the segment 812. The controller may receive asignal indicative of the position of the spacecraft 800, and switch fromdirecting energy and/or propellant from one thruster to another. Forexample, upon detecting the position within the segment 814 b of thespacecraft 800 moving towards the periapsis along the orbit 810, thecontroller may activate the thruster 884 (e.g., a MET thruster) togenerate thrust while maintaining a substantially high specific impulse(Isp). On the other hand, upon detecting the position within the segment814 b of the spacecraft 800 moving towards the periapsis along the orbit810, the controller may activate the thruster 884 (e.g., a direct solarthermal thruster) to generate higher thrust at the cost of a relativelysmaller specific impulse (Isp). In some implementations, the thruster884 may have higher thrust that the thruster 882 and may be configuredto provide thrust throughout the segments 812 and 814 a,b. The thruster882 may be configured for providing thrust for attitude adjustments.

FIG. 9 is a flow diagram of an example method 900 of operating apropulsion system of a spacecraft (e.g., spacecraft 100). At block 910,the method 900 includes determining spacecraft position along atrajectory. Determining the position may include identifying an orbit ofthe spacecraft, identifying a section of the orbit within an ellipticalorbit around a celestial body, and/or identifying spacecraft speed andorientation. One or more sensors (e.g., included in the sensors andcommunications component 120) may communicate the data indicative of thespacecraft position to a flight computer (e.g., the flight computer 150)or another suitable subsystem for configured for computing and controlof subsystems configured for propulsion.

At block 920, the method 900 includes determining based on the spaceposition an operating mode of a thruster system (e.g., the thrustersystem 180). The thruster system may include a first thruster operatingaccording to a first propulsion technique and a second thrusteroperating according to a second propulsion technique. Determining theoperating mode may include determining whether the first thruster and/orthe second thruster are active and providing propulsion to thespacecraft. For example, the first thruster may be a MET thruster andthe second thruster may be a solar thermal thruster. Determining theoperating mode may further include determining a level or amount ofthrust within an operating range of each thruster.

At block 930, the method 900 includes configuring a power system (e.g.,the power system 170) to direct energy according to the determinedoperating mode of the thruster system. The method 900 may includeconverting solar energy to heat and/or electricity. For example, thepower system may include an energy plant configured to convert solarenergy into electricity and to deliver the electricity to the firstthruster. To that end, the energy plant may include photovoltaic cells,a heat engine, and/or a turbine. A controller may cause the power systemto guide solar energy toward the energy plant in response to determiningthat the operating mode of the thruster system includes activating thefirst thruster. The controller may alternatively guide solar energytoward a thermal receiver in a second thruster configured as a solarthermal thruster in response to determining that the operating mode ofthe thruster system includes activating the second thruster.

At block 940, the method 900 includes configuring a propellant system(e.g., the propellant system 190) to direct propellant according to thedetermined operating mode of the thruster system. The propulsion systemmay include a tank storing propellant fluidically coupled to the firstthruster and the second thruster (e.g., as illustrated in FIG. 2). Acontroller may supply the propellant to the first thruster and notsupply the propellant to the second thruster when in a first section ofan orbit, and/or supply the propellant to the second thruster and notsupply the propellant to the first thruster when in a second section ofan orbit, as discussed above in the context of FIG. 8. Furthermore, thecontroller may control an amount of propellant based on the determinedamount of thrust when determining the operating mode includesdetermining the amount of thrust.

In some implementations, directing energy to one of the thrusters or toa power plant includes directing propellant. Thus, at least a portion ofthe method at block 940 may be implemented at block 930. For example,the propulsion system may include a thermal receiver with a first heatexchanger and a second heat exchanger (as illustrated in FIG. 5). Thus,directing propellant also directs the energy that the propellant picksup at one of the heat exchangers.

The following list of aspects reflects a variety of the embodimentsexplicitly contemplated by the present disclosure.

Aspect 1. A propulsion system operating in a spacecraft, the propulsionsystem comprising: a power system configured to collect solar energy; afirst thruster operating according to a first propulsion technique; asecond thruster operating according to a second propulsion technique;and a controller configured to control supply of the solar energycollected by the power system to the first thruster and the secondthruster.

Aspect 2. The propulsion system of aspect 1, wherein: the first thrusteris a microwave electro-thermal (MET) thruster, and the second thrusteris a solar thermal thruster.

Aspect 3. The propulsion system of aspect 1 or 2, further comprising: anenergy plant configured to convert solar energy into electricity and todeliver the electricity to the first thruster; wherein the controllercauses the power system to guide solar energy toward the energy plant.

Aspect 4. The propulsion system of aspect 3, wherein the energy plantincludes a photovoltaic cell.

Aspect 5. The propulsion system of aspect 3, wherein the energy plantincludes at least one of: a heat engine or a turbine.

Aspect 6. The propulsion system of aspect 1 or 2, wherein: the secondthruster includes a thermal receiver, and the controller causes thepower system to guide solar energy toward the thermal receiver.

Aspect 7. The propulsion system of aspect 1 or 2, further comprising athermal receiver including: a first heat exchanger with a first fluidicchannel guiding propellant to generate electric energy for the firstthruster, and a second heat exchanger with a second fluidic channelguiding propellant to the second thruster.

Aspect 8. The propulsion system of aspect 1, wherein the power systemincludes a solar concentrator.

Aspect 9. The propulsion system of aspect 1, wherein: the solarconcentrator is pivotable, and the controller causes the solarconcentrator to pivot toward a first receiver associated with the firstthruster or a second receiver associated with the second thruster.

Aspect 10. The propulsion system of aspect 1, wherein the first thrustercorresponds to a first total amount of thrust and a first specificimpulse, and the second thruster corresponds to a second total amount ofthrust greater than the first total amount of thrust, and a secondspecific impulse smaller than the first specific impulse.

Aspect 11. The propulsion system of aspect 1, wherein the controller isconfigured to: supply the solar energy to the first thruster and notsupply the solar energy to the second thruster when the spacecraft is ina first section of an elliptical orbit around a celestial body; andsupply the solar energy to the second thruster and not supply the solarenergy to the first thruster when the spacecraft is in a second sectionof an elliptical orbit around the celestial body.

Aspect 12. The propulsion system of aspect 1, wherein the controller isconfigured to: supply the propellant to the first thruster and notsupply the propellant to the second thruster when the spacecraft is in afirst orbit; and supply the propellant to the second thruster and notsupply the propellant to the first thruster when the spacecraft is in asecond orbit.

Aspect 13. The propulsion system of aspect 1, further comprising a tankstoring propellant fluidically coupled to the first thruster and thesecond thruster.

Aspect 14. A propulsion system operating in a spacecraft, the propulsionsystem comprising: a tank storing propellant; a first thrusterfluidically coupled to the tank, the first thruster operating accordingto a first propulsion technique; a second thruster fluidically coupledto the tank, the second thruster operating according to a secondpropulsion technique; and a controller configured to control supply ofthe propellant from the tank to the first thruster and the secondthruster.

Aspect 15. The propulsion system of aspect 14, wherein: the firstthruster is a microwave electro-thermal (MET) thruster, and the secondthruster is a solar thermal thruster.

Aspect 16. The propulsion system of aspect 14, wherein the firstthruster corresponds to a first total amount a first range of thrust anda first efficiency of thrust, and the second thruster corresponds to asecond total amount of thrust greater than the first total amount or afirst range of thrust, and a second efficiency of thrust smaller thanthe first efficiency of thrust.

Aspect 17. The propulsion system of aspect 14, wherein the controller isconfigured to: supply the propellant to the first thruster and notsupply the propellant to the second thruster when the spacecraft is in afirst section of an elliptical orbit around a celestial body; and supplythe propellant to the second thruster and not supply the propellant tothe first thruster when the spacecraft is in a second section of anelliptical orbit around the celestial body.

Aspect 18. The propulsion system of aspect 14, wherein the controller isconfigured to: supply the propellant to the first thruster and notsupply the propellant to the second thruster when the spacecraft is in afirst orbit; and supply the propellant to the second thruster and notsupply the propellant to the first thruster when the spacecraft is in asecond orbit.

Aspect 19. The propulsion system of aspect 14, further comprising: afirst valve disposed in a fluid line connecting the tank to the firstthruster; and a second valve disposed in a fluid line connecting thetank to the second thruster; wherein the controller operates the firstand second valves to control the supply of the propellant to the firstand second thrusters.

Aspect 20. The propulsion system of aspect 14, wherein the propellant isat least one of (i) water, (ii) hydrazine, (iii) hydrogen peroxide, or(iii) ammonia.

Aspect 21. A system for storing solar energy in a spacecraft, the systemcomprising: a tank storing propellant; a power system configured to (i)collect solar energy and store at least a portion of the collected solarenergy in the propellant in a vaporized state, and (ii) use a firstportion of the vaporized propellant to generate electric energy; and athruster configured to consume a second portion of the vaporizedpropellant to generate thrust.

Aspect 22. The system of aspect 21, wherein the thruster is a solarthermal thruster.

Aspect 23. The system of aspect 21 or 22, further comprising: amicrowave electro-thermal (MET) thruster, wherein an oscillator uses thegenerated electric energy to generate microwave energy.

Aspect 24. The system of aspect 23, wherein the MET thruster consumesthe propellant to generate thrust.

Aspect 25. The system of aspect 21, wherein the tank is expandable.

Aspect 26. The system of aspect 21, wherein the power system includes asolar concentrator.

Aspect 27. A propulsion system operating in a spacecraft, the propulsionsystem comprising: a tank storing propellant; a fuel cell fluidicallycoupled to the tank and configured to operate using a first portion ofthe propellant as a working fluid; and a thruster fluidically coupled tothe tank and configured to consume a second portion of the propellant togenerate thrust.

Aspect 28. The propulsion system of aspect 27, wherein the propellant isat least one of (i) water, (ii) hydrozene, (iii) hydrogen peroxide, or(iii) ammonia.

Aspect 29. The propulsion system of aspect 27, wherein the thruster is amicrowave electro-thermal (MET) thruster.

Aspect 30. The propulsion system of aspect 27, wherein the thruster is asolar thermal thruster.

Aspect 31. The propulsion system of aspect 27, wherein the secondportion of the propellant includes at least some of the first portion ofthe propellant.

What is claimed:
 1. A system operating in a spacecraft, the systemcomprising: a tank storing a propellant; a thermal receiver configuredto change the propellant from a first phase to a second phase byproviding heat to the propellant; and an energy conversion devicefluidically coupled to the tank and configured to generate electricenergy using the propellant in the second phase.
 2. The system of claim1, wherein the thermal receiver converts solar radiation to heat, andwherein the system further comprises a solar collector providing thesolar radiation to the thermal receiver.
 3. The system of claim 1,wherein the tank includes the thermal receiver, and wherein the tank isinsulated to reduce the loss of propellant heat.
 4. The system of claim3, wherein the tank is expandable.
 5. The system of claim 1, wherein theenergy conversion device is a turbine, and wherein the system furthercomprises a turbine outlet configured to return the propellant to thetank.
 6. The system of claim 1, wherein the energy conversion device isa turbine, and wherein the system further comprises (i) a thruster and(ii) a turbine outlet configured to guide the propellant to thethruster.
 7. The system of claim 1, further comprising a thruster,wherein the energy conversion device supplies electricity for poweringthe thruster.
 8. The system of claim 7, wherein the thruster is amicrowave electrothermal thruster, and wherein the energy conversiondevice supplies electricity for powering a microwave source of themicrowave electrothermal thruster.
 9. The system of claim 1, furthercomprising an electric energy storage device, wherein the energyconversion device charges the electric energy storage device.
 10. Thesystem of claim 1, wherein the propellant is water.
 11. The system ofclaim 1, wherein the first phase is a liquid phase and the second phaseis a vapor phase.
 12. The system of claim 1, further comprising: athruster fluidically coupled to the tank and configured to generatethrust using the propellant in the second phase; and a controllerconfigured to selectively direct the propellant in the second phase toeither the energy conversion device or the thruster.
 13. A methodimplemented in a spacecraft, the method comprising: storing a propellantin a tank; changing the propellant from a first phase to a second phaseby providing heat to the propellant via a thermal receiver; andgenerating, by an energy conversion device fluidically coupled to thetank, electric energy using the propellant in the second phase.
 14. Themethod of claim 13, wherein changing the propellant from the first phaseto the second phase includes converting solar radiation to heat, andwherein the method further comprises providing the solar radiation tothe thermal receiver via a solar collector.
 15. The method of claim 13,wherein the energy conversion device is a turbine, and wherein themethod further comprises returning the propellant to the tank via anoutlet of the turbine.
 16. The method of claim 13, wherein the energyconversion device is a turbine, and wherein the method further comprisesguiding the propellant to a thruster via an outlet of the turbine. 17.The method of claim 13, further comprising supplying, via the energyconversion device, electricity for powering a thruster.
 18. The methodof claim 17, wherein the thruster is a microwave electrothermalthruster, and supplying the electricity for powering the thrusterincludes supplying the electricity for powering a microwave source ofthe microwave electrothermal thruster.
 19. The method of claim 13,further comprising charging, via the energy conversion device, anelectric energy storage device.
 20. The method of claim 13, wherein thefirst phase is a liquid phase and the second phase is a vapor phase. 21.The method of claim 13, further comprising: selectively directing thepropellant in the second phase to either (i) the energy conversiondevice or (ii) a thruster fluidically coupled to the tank and configuredto generate thrust using the propellant in the second phase.